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HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: HQ 0/7 AIRFOIL (hq07-il)
Reynolds number: 200,000
Max Cl/Cd: 36.73 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hq07-il-200000-n5.txt
Download as CSV file: xf-hq07-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 0/7 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6752   0.08429   0.08092   0.0048   1.0000   0.0068
  -8.750  -0.6817   0.07815   0.07483   0.0002   1.0000   0.0067
  -8.500  -0.6931   0.07133   0.06802  -0.0064   1.0000   0.0066
  -8.250  -0.7045   0.06593   0.06255  -0.0092   1.0000   0.0065
  -8.000  -0.7107   0.06025   0.05682  -0.0112   1.0000   0.0064
  -7.750  -0.7133   0.05453   0.05090  -0.0124   1.0000   0.0063
  -7.500  -0.7121   0.04883   0.04492  -0.0127   1.0000   0.0061
  -7.250  -0.7072   0.04310   0.03881  -0.0124   1.0000   0.0059
  -7.000  -0.6990   0.03733   0.03246  -0.0115   1.0000   0.0057
  -6.750  -0.6865   0.03219   0.02676  -0.0103   1.0000   0.0055
  -6.500  -0.6701   0.02795   0.02194  -0.0090   1.0000   0.0054
  -6.250  -0.6503   0.02455   0.01802  -0.0078   1.0000   0.0054
  -6.000  -0.6285   0.02189   0.01492  -0.0067   1.0000   0.0054
  -5.750  -0.6056   0.01976   0.01247  -0.0057   1.0000   0.0056
  -5.500  -0.5825   0.01803   0.01050  -0.0048   1.0000   0.0060
  -5.250  -0.5595   0.01657   0.00884  -0.0039   1.0000   0.0065
  -5.000  -0.5363   0.01532   0.00740  -0.0030   1.0000   0.0078
  -4.750  -0.5127   0.01425   0.00614  -0.0021   1.0000   0.0107
  -4.500  -0.4881   0.01360   0.00542  -0.0016   1.0000   0.0179
  -4.250  -0.4633   0.01313   0.00496  -0.0012   1.0000   0.0293
  -4.000  -0.4386   0.01264   0.00447  -0.0008   1.0000   0.0421
  -3.750  -0.4141   0.01207   0.00387  -0.0003   1.0000   0.0613
  -3.500  -0.3905   0.01137   0.00338   0.0002   1.0000   0.1095
  -3.250  -0.3678   0.01062   0.00302   0.0007   1.0000   0.2025
  -3.000  -0.3456   0.00988   0.00268   0.0013   1.0000   0.3004
  -2.750  -0.3268   0.00888   0.00246   0.0025   1.0000   0.4774
  -2.500  -0.3067   0.00838   0.00242   0.0040   1.0000   0.5906
  -2.250  -0.2853   0.00815   0.00239   0.0055   1.0000   0.6588
  -2.000  -0.2623   0.00798   0.00238   0.0066   0.9980   0.7132
  -1.750  -0.2290   0.00788   0.00228   0.0056   0.9864   0.7532
  -1.500  -0.1942   0.00783   0.00223   0.0042   0.9754   0.7730
  -1.250  -0.1590   0.00779   0.00218   0.0027   0.9636   0.7922
  -1.000  -0.1239   0.00775   0.00215   0.0013   0.9490   0.8112
  -0.750  -0.0907   0.00771   0.00214   0.0005   0.9330   0.8320
  -0.500  -0.0596   0.00768   0.00212   0.0001   0.9179   0.8520
  -0.250  -0.0293   0.00766   0.00211  -0.0001   0.9027   0.8691
   0.000   0.0000   0.00765   0.00211   0.0000   0.8864   0.8865
   0.250   0.0294   0.00766   0.00211   0.0001   0.8692   0.9028
   0.500   0.0596   0.00768   0.00212  -0.0001   0.8521   0.9179
   0.750   0.0907   0.00771   0.00214  -0.0005   0.8320   0.9331
   1.000   0.1239   0.00775   0.00215  -0.0013   0.8112   0.9491
   1.250   0.1590   0.00779   0.00218  -0.0027   0.7923   0.9638
   1.500   0.1942   0.00783   0.00222  -0.0042   0.7730   0.9757
   1.750   0.2290   0.00788   0.00227  -0.0056   0.7531   0.9867
   2.000   0.2623   0.00798   0.00238  -0.0066   0.7130   0.9985
   2.250   0.2847   0.00814   0.00238  -0.0054   0.6592   1.0000
   2.500   0.3061   0.00838   0.00241  -0.0039   0.5921   1.0000
   2.750   0.3262   0.00888   0.00246  -0.0024   0.4784   1.0000
   3.000   0.3452   0.00985   0.00267  -0.0012   0.3047   1.0000
   3.250   0.3673   0.01062   0.00301  -0.0006   0.2027   1.0000
   3.500   0.3901   0.01139   0.00339  -0.0001   0.1079   1.0000
   3.750   0.4138   0.01207   0.00386   0.0004   0.0613   1.0000
   4.000   0.4383   0.01264   0.00447   0.0009   0.0421   1.0000
   4.250   0.4632   0.01311   0.00493   0.0012   0.0286   1.0000
   4.500   0.4881   0.01361   0.00541   0.0016   0.0179   1.0000
   4.750   0.5126   0.01429   0.00618   0.0021   0.0103   1.0000
   5.000   0.5364   0.01531   0.00739   0.0029   0.0079   1.0000
   5.250   0.5597   0.01657   0.00884   0.0038   0.0066   1.0000
   5.500   0.5828   0.01803   0.01050   0.0047   0.0060   1.0000
   5.750   0.6059   0.01978   0.01249   0.0057   0.0056   1.0000
   6.000   0.6288   0.02191   0.01494   0.0066   0.0054   1.0000
   6.250   0.6508   0.02457   0.01804   0.0077   0.0054   1.0000
   6.500   0.6706   0.02799   0.02199   0.0089   0.0054   1.0000
   6.750   0.6871   0.03224   0.02681   0.0101   0.0055   1.0000
   7.000   0.6995   0.03742   0.03256   0.0114   0.0057   1.0000
   7.250   0.7080   0.04313   0.03874   0.0122   0.0059   1.0000
   7.500   0.7126   0.04900   0.04510   0.0125   0.0061   1.0000
   7.750   0.7142   0.05461   0.05099   0.0121   0.0063   1.0000
   8.000   0.7118   0.06029   0.05687   0.0110   0.0064   1.0000
   8.250   0.7053   0.06608   0.06270   0.0089   0.0065   1.0000
   8.500   0.6936   0.07156   0.06826   0.0060   0.0066   1.0000
   8.750   0.6824   0.07844   0.07511  -0.0006   0.0067   1.0000
   9.000   0.6763   0.08449   0.08112  -0.0052   0.0068   1.0000
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